Gas turbine engines



Dec. 31, 1968 D. RICH 3,418,808

GAS TURBINE ENGINES Filed Jan. 5, 1967 Sheet of2 Dec. 31, 1968 D. RICHGAS TURBINE ENGINES [v vs r02. D0 v/b R/cw m a m/ Sheet Filed Jan. 5,1967 lira. 6. I8 30b firroeuszs.

United States Patent 3,418,808 GAS TURBINE ENGINES David Rich, 3741 E.Pasadena Ave., Phoenix, Am. 85018 Continuation-impart of applicationSer. No. 562,806, July 5, 1966. This application Jan. 3, 1967, Ser. No.606,777

6 Claims. (Cl. 60--226) ABSTRACT OF THE DISCLOSURE This improvement ingas turbine engines incorporates special internal cooling ducts whichprovide air passageways adjacent to the internal components of theengine. These cooling ducts are concentric with the combustion duct of agas turbine engine and function as by-pass pasageways which provide alarge stream of cooling air inside the engine and through the turbine.

This application is a continuationdn part of my copending US. patentapplication, Ser. No. 562,806, filed July, 5, 1966.

In gas turbine engines an important limitation to the power which can beproduced and the efiiciency of the engine is the amount of heat theinternal components such as the turbine wheels, turbine blades, fixedvanes and engine bearings can Withstand before they will fail. Varioustechniques have been used to alleviate the heat problem of the internalengine components. These techniques include using materials which canwithstand higher temperatures, high temperature lubricants, hollow aircooled turbine blades and bleeding cooling air from various stages ofthe compressor to cool the hot inner parts of the engine. However,limitations to the performance of gas turbine engines due to interiorengine heat still exist despite these techniques, which often areinefficient and wasteful of energy that otherwise could be developed bythe engine.

Hence, it is a primary object of this invention to provide a novelstructure for cooling the internal components of a gas turbine enginewhich is superior to prior techniques.

It is a further object of this invention to provide a cooling structurefor gas turbine engines which is efficient and does not waste thepotential energy of the engine.

A still further object of this invention is to provide a by-pass typeair diversion cooling structure in a turbine engine which is especiallyeffective in cooling the internal engine components.

Other objects and advantages will be apparent from the followingdescription and drawings wherein:

FIGURE 1 is a simplified longitudinally sectioned schematic view of asingle spool by-pass type gas turbine engine incorporating the uniqueinternal by-pass cooling system which is the subject of this invention;

FIGURE 2 is a cross-sectional view taken along the line 2-2 in FIGURE 1showing the diversion of the inlet air into the engine;

FIGURE 3 is a cross-sectional view taken along the line 33 in FIGURE 1showing the air passages through and around the high pressure stages ofthe axial compresor;

FIGURE 4 is a fragmented schematic view of the forward portion of aturboshaft type gas turbine engine embodying an internal by-pass coolingpasageway in accordance with this invention;

FIGURE 5 is a cross-sectional view taken along the line 55 in FIGURE 1;

3,418,808 Patented Dec. 31, 1968 FIGURE 6 is a longitudinally sectionedschematic view of a modified form of the gas turbine engine;

FIGURE 7 is a cross-sectional view taken along the line 7-7 in FIGURE 6;and

FIGURE 8 is a fragmented schematic view of the board portion of aturboshaft type gas turbine engine embodying the modified form of thisinvention as in FIGURES 6 and 7.

Though applicable to all other types of gas turbine engines referencewill be made herein primarily to aircraft by-pass type gas turbineengines as schematically illustrated in the form modified in accordancewith this invention in FIGURES 1, 2, 3 and 5, the conventional parts ofwhich will be described only briefly. In this type of engine when inoperation air enters through a circular front inlet 10' in the forwardend of a hollow cylindrical external engine casing 12 that surrounds theengine components. The air then passes through a plurality of staticguide vanes 14 which are secured at their inner ends to and radiate froma rounded centrally positioned nose cap 16 and which are secured attheir outer ends to the inside of engine casing 12 just to the rear ofits inlet 10.

After passing through the static guide vanes 14 and concentricallyaround nose cap 16 the air then passes through a low pressure axialcompressor 18 consisting of one or more stages concentrically disposedaround and radiating from the forward part of a rotatably mountedelongated cylindrical drive shaft housing 20 axially extending, withincasing 12 and within which is axially mounted an elongated circularcylindrical drive shaft 21. Thereafter, the air is diverted throughthree major concentric passageways the outermost of which is theexternal by-pass duct 22, the intermediate one being a combustion duct24 leading through a high pressure compressor to the radially positionedcombustion chambers 26, and in terior of combustion duct 24, an internalbypass duct 28. Additionally, a small relative portion of the airpassing through the intermediate combustion duct 24 is diverted so as toby-pass internally around the combustion chambers 26 through aconcentric turbine cooling by-pass passageway 24a. The internalcomponents of the engines to be 'briefiy referred to when relevant tothis invention are supported within the external engine casing in aconventional manner. The rotating parts are mounted on suitable bearingslubricated by conventional systems.

The outside periphery of the external by-pass duct 22 is defined by thecylindrical inner wall 12a of the external engine casing 12 from therear of the low pressure compressor 18 to the rear outlet 29 of theengine. Its inner boundary is defined by the outside surface 30a of aconcentric annular outer combustion duct casing 30 secured to the innerwall 12a of casing 12 by radiating support arms 31 and extending fromits annular inlet lip 30b disposed at the rear of low pressurecompressor 18 to the front of the exterior combustion chamber wall 32,the outer surface 32a of the exterior combustion chamber wall 32 and theouter surface 34a of an annular outer exhaust casing 34 extending fromthe rear of the exterior combustion chamber wall 32 to the rear outlet29 of the engine. Thus, air diverted through the external by-pass duct22 is discharged from the engine through the concentric external by-passexhaust 36 in the rear outlet 29 of the engine, picking up heat energyand consequent velocity on the way so as to augment the thrust of theengine.

The exterior of the combustion duct 24 is defined by the inside surface300 of the annular outer combustion duct casing 30. The interior of thecombustion duct 24 is defined from front to rear by the outer surface38a of a concentric annular inner combustion casing 38 supported by theradiating arms 39 connected to outer combustion duct casing 30, theouter surface 40a of an annular compressor wall 40 and the insidesurface 42a of the combustion chamber lip 42 of an interior combustionchamber wall 44, the latter which, together with the external combustionchamber walls 32 form combustion chambers 26. Air passing through thiscombustion duct 24 is compressed further by a high pressure compressor46, the blades of which are in duct 24, before it enters the combustionchambers 26 into which fuel is normally sprayed by a fuel injector 48from a pressurized source (not shown). Upon ignition of the fuel-airmixture in a conventional manner the gases are discharged fromcombustion chambers 26 so as to provide thrust from the engine through acombustion exhaust passageway 50, the exterior of which is defined bythe interior surface 34b of the annular exhaust casing 34. The interiorof exhaust passageway 50 is defined, from front to rear, by the exteriorsurface 52a of an annular turbine casing 52 and the exterior surface 54aof an annular truncated tail cone 54. As these gases pass throughexhaust passageway 50 they rotate the turbine 56 the blades-of which aredisposed in passageway 50 and which is connected by radially extendingturbine arms 58 to drive shaft housing 20. This, in turn, rotates thelow pressure compressor 18 and the high pressure compressor 46, thelatter being connected to the common drive shaft housing 20 by radiatingcompressor arms 60.

The description of the illustrated aircraft gas turbine engine has beenconventional to this point save for the addition of passageway 24a andduct 28 and is only intended as an orientation to the improvementthereto hereinafter described. That is, a unique improvement in gasturbine engines in accordance with my invention is to provide such anengine with an internal by-pass duct 28 which not only provides theadvantages in propulsion of the normal by-pass type jet engine, butcools the internal components of the engine so that it may operate evenmore efficiently. A further improvement herein is the turbine coolingby-pass passageway 24a which opens concentrically from the combustionduct 24 at the rear of high pressure compressor 46 and by-passes thecombustion chambers 26 so as to direct a blast of cooling air onto thebase of the blades of the turbine 56 immediately to the rear of thecombustion chambers 26. Thus, the cooling effect of this divertedairflow allows the turbine and interior engine components to operate ata higher combustion temperature with consequent greater efficiency. Alsothe air traveling through the by-pass system not only will significantlycool the engine interior but is not wasted as is bled air, since theheat that it absorbs will increase its exhaust velocity therebyaugmenting the thrust of the engine.

The interior of this internal by-pass duct 28 is defined by the exteriorsurface 20a of the drive shaft housing 20 from just to the rear of thelow pressure compressor 18 to the rear end 20b of the housing 20. Fromthere to the rear outlet 29 of the engine the interior of the internalby-pass duct 28 is defined by the exterior surface 62a of a cone-shapedexhaust bullet 62 centrally mounted adjacent the outlet 29 within casing12 and within which the rear end 21b of the axial shaft 21 is journaledfor rotatable movement the front end 21a being rotatably journaled innose cap 16. The outer periphery of the internal bypass duct 28 isdefined, from front to rear, by the inner surface 38b of the annularinner combustion casing 38, and then the continuous inner surface 40b ofthe compressor casing 40, the inner surface b of a combustion chamberby-pass wall 45, the inner surface 52b of the annular turbine casing 52,and the inner surface 54b of the annular truncated tail cone 54. Hence,air diverted through the internal by-pass duct 28 will be dischargedfrom the engine through a concentric internal by-pass exhaust 64surrounding bullet 62 and will pass a large mass of cooling air close tothe bulk of the internal components of the engine in a substantiallyuninterrupted concentric axial fiow therethrough. The massive coolingeffect produced thereby not only allows increased temperature in thecombustion chamber but permits the more practical inclusion of variousheat sensitive accessories within the engine which ordinarily would haveto be mounted outside the engine. Also, since the engine exhaust issandwiched between two by-pass passageways its noise will beacoustically muffled to a great extent.

It can be seen that the air diverted into internal by-pass duct 28performs its cooling function in a distinctly different manner frommerely bleeding air into the engine interior as was done prior to thisinvention. For instance, whereas air bled into the engine interiorgenerally reduces the engine efiiciency since it, in effect, is wastedso far as any direct thrust is produced thereby, air which fiows axiallyin by-pass duct 28 adds to the thrust of the engine in accordance withnormal by-pass principles. The ratio of air b y-passed to that used forcombustion varies according to performance requirements which controlthe specific size and shape of each of the air passageways.

The relatively small concentric turbine cooling by-pass passageway 24ais defined externally by an internal surface 42b of the lip 42 and aninternal surface 44b of the interior combustion chamber wall 44.Internally the turbine cooling by-pass passageway 24a is defined by theexternal surface 45a of the combustion chamber by-pass wall 45 whichextends continuously between compressor casing 40 and turbine case 52.By-pass wall 45 is spaced slightly inwardly from combustion chamber wall44 so that the entry into cooling by-pass passageway 24a from combustionduct 24 is adjacent to forward lip 42 and the exit therefrom is adjacentto the rear of combustion chambers 26 and the forward portion of theblades of turbine 56. Thus, without any leakage the entire amount of aircompressed by the compressor 46 is discharged through the turbine 56.That is, a major portion travels through the combustion chambers 26 toprovide thrust for the engine and a minor portion is diverted around thecombustion chambers 26 through by-pass passageway 24a to cool the firststage turbine blades and reduce the blade temperatures to a safe level.

The internal by-pass cooling system in accordance with my invention canbe used with tur-boshaft type gas turbine engines as shown in FIGURE 4.In this schematic drawing the rear portion of the engine is essentiallythe same as described relative to FIGURE 1 and so is not shown nor itsdescription reiterated here. In the fragmented front portion of theengine shown the same numerals are used to designate like parts andprimed numerals used to designate modified parts.

Thus, the turboshaft aircraft engine in FIGURE 4 has a cylindricalexternal engine casing 12 with a circular forward inlet lip 10.concentrically disposed within engine casing 12 are an external by-passduct 22, a combustion duct 24, a turbine cooling by-pass passageway 24aand an internal by-pass duct 28 surrounding a cylindrical drive shafthousing 20. Axially mounted within the drive shaft housing 20 is a driveshaft 21 which is journaled for rotatable movement and passes forwardthrough nose cap 16. In turn, to its forward end may be attached apropeller or other power take-off device (not shown) by suitablemodifications.

As the previously described form of this invention air will enter theengine through inlet 10, pass through radial guide vanes 14 and througha low pressure compressor 18' shown as a single stage fan type here.Thereafter, the air will be diverted through each of the three ducts 22,24, 28 and by-pass passageway 24a and eventually be discharged aspreviously described. The passageway through duct 28 allows a large massof air at low pressure to flow therethrough adjacent to the internalcomponents of the engine thereby producing a massive cooling effect witha minimum of power loss. And, as in the previously described embodimentof this invention, a relatively small amount of air from the highpressure compressor 46 is diverted around the combustion chambers 26 todirect a blast of cooling air into the turbine blades disposedrearwardly of the combustion chambers 26.

The modified form of this invention shown in FIG- URES 6 and 7 issimilar to that previously disclosed, with like numbers designating likeparts, save that it incorporates additionally a small concentricexterior combustion chamber by-pass passageway 33 which receivescompressed air from the high pressure compressor 46 and diverts acooling blast of air around the outside of the combustion chambers 26onto the tips of the blades of turbine 56 immediately to the rearthereof. Thus, the cooling effect of the engine is further increased bythis passageway 33 without loss of efficiency since this bypassed air isstill diverted essentially in an axial flow and precisely onto the partsof the turbine 56 where it will do the most good.

An annular interior combustion chamber wall 35 spaced inwardly from theexterior combustion chamber wall 32 within combustion chambers 26 formsthe interior of passageway 33 with its outer surface 35a. The exteriorof passageway 33 is defined by the inner surface 32b of the exteriorcombustion chamber wall 32.

The efliciency of a t-urboshaft type gas turbine engine is alsoincreased by the addition of an exterior combustion chamber by-passpassageway 33 constructed in a manner substantially the same aspreviously described. Such a passageway 33 is shown in FIGURE 8 withlike parts being designated by like numerals.

The amount of air that bypassses the combustion chamber and is appliedto cool the turbine blades could be controlled so that when maximumpower is required and the maximum blade cooling is essential, the amountof air bypassing the combustion chamber is at a maximum. However, whenthe engine power is reduced, this volume of bypassed air could bereduced or cut off completely, and the air could be reduced or cut offcompletely, and the cooling effect of the internal air stream would beadequate to keep the turbine running at a safe temperature.

The foregoing description of this invention has been illustrated inconnection with aircraft bypass gas turbines as preferred embodiments.However, the advantages described can also be attained with enginesusing centrifugal and mixed axial and centrifugal compressors. Also, theinvention can be used in stationary engines where the exhaust thrust isnot used for propulsion and the maximum shaft power is needed. When usedfor these purposes modifications in the structure would normally includeelimination of the external duct 22 and reduction of the relative sizeof the internal by-pass duct 28 to that required for cooling only. Also,though the engine illustrated is of the single spool type the inventioncan be used with multiple spool engines.

Although I have herein shown and described my invention in what I haveconceived to be the most practical and preferred embodiment, it isrecognized that departures may be made therefrom within the scope of myinvention, which is not to be limited to the details disclosed hereinbut is to be accorded the full scope of the claims so as to embrace anyand all equivalent structures and devices.

I claim:

1. In a gas turbine engine comprising: a hollow cylindrical externalengine casing embodying an intake opening in its front end and anexhaust opening in its rear end; an axially positioned nose cap mountedwithin said engine casing adjacent to said intake opening therein; guidevanes radiating from said nose cap as a support therefor, said guidevanes being secured to the inner surface of said engine casing; a tailcone bullet axially positioned within said engine casing supportedtherein adjacent to said exhaust opening in the rear of said casing; anelongated drive shaft axially extending within said external enginecasing, the front end of which is rotatably supported by said nose capand the rear end of which is rotatably supported by said tail conebullet; a cylindrical drive shaft housing concentrically supported bysaid drive shaft and substantially encompassing said drive shaft betweensaid nose cap and said tail cone bullet; a low pressure axial compressorradiating from said drive shaft housing immediately to the rear of saidguide vanes; a high pressure axial compressor within said engine casingimmediately to the rear of said low pressure axial compressor connectedby radiating compressor arms to said drive shaft housing; at least oneradially positioned combustion chamber supported by and within saidengine casing immediately to the rear of said high pressure compressor;means for injecting fuel into said combustion chamber; an axiallyrotatable turbine immediately to the rear of said combustion chamberwithin said engine casing connected to said drive shaft housing byradiating turbine arms, the combination with: a concentric combustionduct defining a passageway whereby air is diverted from the rear of saidlow pressure compressor through said high pressure compressor into saidcombustion chamber; a concentric combustion exhaust duct to the rear ofsaid combustion chamber defining a passage-way from the rear of saidcombustion chamber through said turbine to the exhaust opening in saidengine casing; a concentric external non-bifurcated by-pass ductdefining a passageway'within said engine casing from the rear of saidlow pressure compressor exterior of the high pressure compressor, thecombustion chamber, the turbine and exhausting exterior of saidcombustion exhaust; and a concentric internal non-bifurcated by-passduct defining a passageway within said engine casing from the rear ofsaid low pressure compressor, interior of the high pressure compressor,the combustion chambers, the turbine and exhausting interior of saidcombustion exhaust, said internal by-pass duct surrounding said driveshaft housing so that air flowing therethrough will cool the hotinternal engine components.

2. A gas turbine engine as defined in claim 1 which includes aconcentric turbine cooling by-pass passageway having an inlet openingfrom the rear of said high pressure compressor adjacent to the entryinto the combustion chamber, an air flow conduit passing around saidcombustion chamber and an outlet opening into a combustion exhaust ductto the rear of said combustion chamber thereby diverting a relativelysmall portion of compressed cool air onto the turbine.

3. A gas turbine engine as defined in claim 1 which includes a pair ofconcentric passageways which both open from said high pressurecompressor to divert air generally axially around the combustionchamber, one of said passageways being formed Within the engineinternally of said combustion chamber and providing an air exit adjacentthe base of the turbine blades to the rear thereof and the other of saidpassageways being formed within the engine externally of said combustionchamber and providing an air exit adjacent the tips of said turbineblades.

4. A gas turbine engine comprising: a hollow cylindrical external enginecasing embodying an intake opening in its front end and an exhaustopening in its rear end; an axially extending drive shaft housingrotatably mounted within said external engine casing; an axialcompressor including a forward low pressure portion and a rear highpressure portion radiating from said drive shaft housing to the rear ofsaid intake opening in said engine casing; a turbine radiating from saiddrive shaft housing forward of said exhaust opening in the rear of saidengine casing; a combustion chamber supported in said engine casingspaced from said drive shaft housing and positioned between said axialcompressor and said turbine; a combustion passageway within said enginecasing extending from the rear of the high pressure portion of saidcompressor into said combustion chamber and exhausts from said chamberthrough said turbine to said exhaust opening of said casing; and aconcentric axial, non-bifurcated internal by-pass duct surrounding saiddrive shaft housing and extending from the rear of the low-pressureportion of said compressor to said exhaust opening of said casing andby-passing said high pressure portion of said compressor, said internalby-pass d-uct defining a space between said combustion passageway andsaid drive shaft housing, said internal bypass duct providing asubstantially uninterrupted concentric axial conduit about said driveshaft housing for the flow of air through said engine whereby in use alarge quantity of air passing said low pressure portion of saidcompressor is axially urged through the by-pass conduit, and will coolthe internal components of said engine.

5. A gas turbine engine as defined in claim 4 which includes aconcentric turbine cooling by-pass passageway which opens from the rearof the high pressure portion of the compressor adjacent to the inlet ofthe combustion chamber and defines a continuous air duct exterior ofsaid combustion chamber with an outlet opening into the turbine, saidby-pass passageway being relatively small compared to the combustionchamber thereby diverting a small amount of compressed, relatively coolair onto said turbine.

6. A gas turbine engine as defined in claim 4 which includes a pair ofconcentric passageways which both open from and are in direct line withsaid high pressure compressor to divert air generally axially around thecombustion chamber, one of said passageways being formed within theengine internally of said combustion chamber and providing an air exitadjacent the base of the turbine blades to the rear thereof and theother of said passageways being formed :within the engine externally ofsaid combustion chamber and providing an air exit adjacent the tips ofsaid turbine blades.

References Cited UNITED STATES PATENTS 2,409,176 10/1946 Allen 60-2622,930,190 3/1960 Rogers 60226 2,933,886 4/1960 Sharma 60-262 FOREIGNPATENTS 1,297,052 5/ 1962 France.

JULIUS E. WEST, Primary Examiner.

DOUGLAS HA'RT, Assistant Examiner.

US. Cl. X.R.

